This invention features an open-lattice, foldable, self-deployable structure which can be used for solar sails, heat shields, communication devices, sensors and sensor arrays, power generation devices, and the like.
Ultra-light weight, large spacecraft structures can be used for a variety of purposes including solar sails, heat shields, communication devices, power generation devices, and the like. The structures must be compatible for deployment, self deployable, and upon deployment, sufficiently rigid and stable and unfolding to a shape exactly the same as it was before it was folded compactly for deployment in space. In the prior art, mechanical systems and inflatable systems were used.
Mechanical deployment systems are among the best understood of prior art spacecraft mechanisms. See U.S. Pat. No. 4,579,302 incorporated herein by this reference. Virtually all spacecraft mechanically deploy at least one component to gather solar power, radiate heat, or act as an antenna. Because of their mission critical nature, these systems have been designed for extreme reliability with margins of safety to accommodate variations in manufacture, changes in lubrication states, and other factors.
Mechanical systems, however, have both theoretical and practical limitations that make it difficult to apply them to ultra-low density requirements of lightly loaded structures. In a mechanical deployment system, the mechanisms that provide the deployment become parasitic mass once the deployment is complete. The hinge or latch mechanisms of these mechanical systems are also generally structurally inefficient and there are practical limitations to how small such a mechanism can reasonably be fabricated and still be reliable. Moreover, the parasitic mass added by the hinges, latches, and actuators reduces overall system performance, limiting the acceleration of solar sails, for example. Finally, the coiled longeron and cable tension systems currently developed and representative of the state of the art in mechanical packaging density generally attain, at best, a length-to-packing ratio of approximately 20 to 1. For a 500 meter boom, this means that the stowed length of the structure would approach 25 meters, exceeding the payload envelope of all existing launch systems.
For several types of spacecraft structures, the loads on the structure are extremely small once the spacecraft is in its final orbit. These forces, typically 2 to 20 lbs., are typically far below the load limit of the structure deployed solar array, sensor, antenna, etc. This means that the structure is more massive than it theoretically has to be, thus increasing the costs of delivering the spacecraft. To reduce the launch mass and thus launch cost of future spacecraft, recent research has proposed to develop xe2x80x9cGossamerxe2x80x9d spacecraft that take advantage of the very light loading environment of space and reduce structural mass and thus launch costs.
To meet the needs of Gossamer spacecraft missions and to rectify the shortcomings of mechanical deployment systems, various government and commercial organizations have been investigating the use of inflatable membranes as structural elements. See, for example, U.S. Pat. No. 3,477,622.
These systems use a thin film membrane that is pressurized in space by either a compressed or chemically stored gas. Once the desired shape has been achieved, the membrane supports the structural loads and either the original inflatant gas or a secondary rigidization system holds the structure in shape. Although such inflatable systems offer the potential of very low mass and very high packaging efficiency for missions that have low ultimate load requirements, there are again both theoretical and practical limitations that have thus far proven to be significant limitations to the use of inflatable systems as compressive structural elements.
In order for a columnar member to be inflated, it must include a continuous pressure vessel so that it will contain the inflatant gas. But, in order for it to be a mass efficient compression column, it must have extremely thin walls. The combination of these requirements leads to very thin membranes in which surface imperfections dominate the local structural behavior. Prior research has shown that the effects of even minor localized deviations from the ideal shape can greatly reduce the stiffness and strength of the resultant column.
Moreover, inflatable structures suffer from the limitation that low level leaks and micro-meteoroid punctures can eventually drain the gas to the point that the structures can no longer hold their shape.
In order to overcome the problems associated with gas leaks and punctures, aluminum based films have been used requiring very large quantities of gas to provide the internal pressure necessary to yield the wall material even when it is very thin. Moreover, the gas used to provide this internal pressure adds to the non-structural launch mass of the system. Compounding this mass problem is the issue that it is difficult to exactly control the yield point of many thousands of square feet of thin film material. As some parts of the film yield past their desired shape, others haven""t started to yield yet. The result is an imperfect shape, inadequate performance, and reduced mission reliability.
To avoid these problems, several developers of light weight inflatable columns have developed techniques to rigidized the inflated shape using resin systems that are cured to hard shape in orbit, once the deployment is complete. The limitations of these systems center on the mass and reliability penalties associated with doing complex material processing in the remote, zero-g space environment. Since the resin system must be consolidated and cured in orbit, a variety of challenges must be overcome. First, the resin must maintain its character for as long as two years of storage and withstand the high temperatures of the launch pad. Second, the resin must be consistently and reliably distributed throughout the column. The third requirement is that there must be some means to cure the epoxy which must also be light weight and reliable. Resins cured by ultraviolet light provided by the sun have low glass transition temperatures and thus are prone to strength loss at the operating temperatures and degradation after continued exposure. Solvent loss based systems have a fairly large mass penalty associated with them since large amounts of non-structural solvent must be carried as part of the mass launch. Also, in orbit, thermal cure systems require substantial amounts of energy and require the additional, non-structural mass of resistive heaters and insulation.
Finally, these techniques suffer from the limited ability to verify their function on the ground. Since the rigidization of a material, either by yielding it or by solidifying a resin system is generally a one-time process, extreme and expensive methods must be undertaken to ensure that the process will proceed as planned.
The advantage of open lattice isogrid tubular structures are known, but such prior art structures are made of rigid members and thus they cannot be folded into a compact package for space deployment. See Mikulas Jr., Martin, M., NASA Technical Memorandum 78687, Structural Efficiency of Long Lightly Loaded Truss and Isogrid Columns for Space Applications (July 1978). U.S. Pat. No. 5,977,932, incorporated herein by this reference, reports studies concerning self-deploying helical structure antennas made of helical strips and rings. This structure has poor structural qualities when compared to isogrid structures and thus requires mechanical booms (see FIGS. 4-17) for mechanical strength.
It is therefore an object of this invention to provide a foldable, self-deployable open lattice structure.
It is a further object of this invention to provide such a structure which is reliable.
It is a further object of this invention to provide such a structure which does not require any hinge, latch, or actuator mechanisms which would add parasitic mass to the structure.
It is a further object of this invention to provide such a structure which can be manufactured at a cost substantially less than structures including expensive and complex hinge, latch, and actuator mechanisms.
It is a further object of this invention to provide such a structure capable of a length-to-packing ratio much greater than prior art mechanical systems.
It is a further object of this invention to provide such a structure which overcomes the problems associated with imperfections in prior art inflatable systems.
It is a further object of this invention to provide such a structure which maintains the proper shape upon deployment.
It is a further object of this invention to provide such a structure which is not susceptible to leaks or micro-meteoroid punctures.
It is a further object of this invention to provide such a structure which does not require an inflating gas for deployment.
The invention results from the realization that, contrary to conventional wisdom, structural members made of composite materials can be designed to bend and even twist to a point below the material""s yield point and that the advantages of open-lattice structures can be fully realized and made foldable for compact storage and made self-deployable if the longitudinal composite members are made bendable and if the diagonal composite members are made both bendable and twistable.
Once such a structure is fabricated, it can be folded or rolled into a compact shape for storage during deployment and then once in position in space released whereupon the structure resurrects itself automatically to the exact shape it was before folding to deploy solar sails, heat shields, communication devices, power generation devices, and the like.
This invention features an open-lattice, foldable, self-deployable structure comprising a plurality of spaced, bendable and twistable diagonal members disposed to span between and intercept longitudinal members thereby forming cells, each cell being bounded by a portion of two spaced longitudinal members and two spaced diagonal members; and means for joining the diagonal members to the longitudinal members at the cell boundary intersections of the diagonal members and the longitudinal members. The longitudinal members are made of a material which bends by a predetermined amount below the material""s yield point and the diagonal members are made of a material which both bends and twists by a predetermined amount below the material""s yield point so that the structure can be collapsed and then self-resurrected.
The longitudinal members and the diagonal members are typically made of a composite material including fibers in a resin matrix. The resin matrix may be thermoplastic material. The cells many have the shape of isosceles triangles when the structure is in the shape of a cylinder having a longitudinal axis, the longitudinal members typically extending along the direction of the longitudinal axis and the diagonal members include one set which extend around the cylinder in a first direction and another set which extend around the cylinder in a second direction.
The longitudinal members are typically continuous from one end of the cylinder to the other. The diagonal members are also typically continuous from one end of the cylinder to the other. The cells may be triangular in shape, the apexes of each cell each including a portion of two diagonal members intersecting the longitudinal members.
The means for joining the diagonal members may include an adhesive securing the diagonal members to the longitudinal members. If the diagonal members and the longitudinal members are made of a composite material including fibers, the means for joining preferably includes the fibers of a longitudinal member inter-woven with the fibers of a diagonal member. A resin secures the woven fibers at the joints between the longitudinal members and the diagonal members.
The open lattice, foldable, self deployable structure of this invention may be embodied in structural shapes other than columns including trusses and other three-dimensional lattice structures.